1. Field of the Invention
The present invention relates to a turbine blade of a gas turbine or the like and a gas turbine equipment using this turbine blade.
2. Description of the Prior Art
FIG. 5 is a schematic explanatory view of a structure of a turbine portion and a cooling air system for cooling this turbine portion in a gas turbine equipment in the prior art.
The turbine portion comprises a rotational portion of a rotor 1 and a turbine moving blade 2 and a stationary portion 5 of a casing 3, a turbine stationary blade 4, various supporting members and the like.
In the turbine portion, a high temperature high pressure combustion gas supplied from a combustor 6 is converted into a high velocity flow by the turbine stationary blade 4 to rotate the turbine moving blade 2 for generation of power.
Construction members of the rotational portion and the stationary portion which are adjacent to the combustion gas need to be cooled so that their temperature due to heat input from the combustion gas may not exceed their respective allowable temperature and, for cooling of the rotational portion having the rotor 1 and the turbine moving blade 2, it is usual that cooling medium 7 is supplied as shown by arrows in FIG. 5.
The cooling medium 7 is often a bleed air or discharge air taken from a compressor (not shown) or sometimes the bleed air or discharge air once supplied into a cooler (not shown) and cooled to an appropriate temperature.
Further, as the cooling medium to cool the mentioned portions, there is recently a case where steam from an outside system is applied in place of the bleed air or discharge air from the compressor, but herebelow description will be made based on the cooling air system which is generally employed as a typical example.
While the cooling medium 7 flowing in the rotational portion takes a route to flow through an interior of the rotor 1 to enter an interior of the turbine moving blade 2 for cooling thereof and then to join into a combustion gas path, in the case of using steam as the cooling medium as mentioned above, the cooling medium which has been heat-exchanged by cooling the turbine moving blade 2 and the like is recovered so that thermal energy thereof may be made use of in an outside system and thermal efficiency of the plant may be enhanced.
In the gas turbine equipment having the mentioned basic structure, description will be made concretely on the prior art turbine portion thereof with reference to FIGS. 6 to 10.
FIG. 6 is a longitudinal cross sectional view showing a main structure of a prior art turbine moving blade, FIG. 7 is a perspective view showing a main structure of a prior art turbine stationary blade, FIG. 8 is an enlarged view of a part of the turbine stationary blade of FIG. 7, FIG. 9 is a qualitative explanatory view showing a metal temperature behavior due to thickness difference between thickness of a turbine moving blade trailing edge portion and that of a platform in the prior art, and FIG. 10 is likewise a qualitative explanatory view showing a metal temperature behavior due to thickness difference between thickness of a turbine stationary blade trailing edge portion and that of a shroud in the prior art.
In a leading edge portion of the turbine moving blade 2 which is exposed to an especially high temperature combustion gas, in order to stand a high thermal load, it is usual to provide a cooling passage 8 through which the cooling medium 7 is supplied for effecting a convection cooling in the turbine moving blade 2.
Cooling passage in the moving blade is often constructed to repeat several turnings so as to form a serpentine passage on design demand, wherein the passage turns at a turning portion 11 provided in the vicinity of a tip portion 9 of the turbine moving blade 2 and a joint portion 10 of the turbine moving blade 2.
Thus, the cooling medium 7 flows through the cooling passages to cool the interior of the turbine moving blade 2. However, in case the turbine moving blade 2 is one which receives higher thermal load, there is provided a film cooling hole 12 in a blade surface of the turbine moving blade 2 and a portion of the cooling medium 7 is blown therethrough onto the blade surface on the combustion gas path side so that the blade surface may be covered by a low temperature air curtain and thereby a film cooling for reducing the thermal load from the blade surface as well can be effected.
On the other hand, a trailing edge portion 14 of the turbine moving blade 2 is usually designed to be relatively thin in order to reduce an aerodynamic loss of the combustion gas and, for this purpose, if the turbine moving blade 2 is to be cooled, a pin fin cooling or a slot cooling by way of many slots is employed for cooling the interior of the blade, or the film cooling by way of blowing air from a ventral side surface of the blade through the film cooling hole is effected.
In case of the turbine stationary blade 16, in order to form a gas flow path, structure of the blade is made such that an inner end of a blade profile portion 17 is inserted into an inner shroud 18 and an outer end of the blade profile portion 17 is inserted into an outer shroud 19, and while this set of one inner shroud 18 and one outer shroud 19 is usually provided for each of the turbine stationary blades 16, there is also such a case where the set of one inner shroud 18 and one outer shroud 19 is provided so as to cover a plurality of the turbine stationary blades 16.
The turbine stationary blade 16 is usually formed by precision casting and is then worked by machining, wherein the inner shroud 18, the outer shroud 19 and the blade profile portion 17 are generally formed integrally by casting.
As mentioned above, the platform 15 supporting the turbine moving blade 2 forms a part of the gas flow path in an axial flow turbine and is made relatively thicker as compared with the trailing edge portion 14 of the blade so as to stand centrifugal force or the like.
For this reason, in operation of the gas turbine including start and stop, load change or the like, there may arise an excessively large temperature difference between the platform 15 and the blade trailing edge portion 14, by which thermal stress is liable to occur at a transition time or in a steady operation time so that there is a risk to cause cracks and if the cracks occur, there is a problem to damage a reliability of the turbine moving blade.
Also, in the turbine stationary blade 16, in order to reduce an aerodynamic loss, a trailing edge portion 20 of the blade is designed as thin as possible and, on the other hand, the inner shroud 18 and the outer shroud 19 are usually designed relatively thicker for holding the strength. Thus, like the turbine moving blade 2, there is a problem that cracks are considered to occur by the thermal stress following a start and stop of the gas turbine or the like, which results in damaging the reliability.
The mentioned relation between the moving blade trailing edge portion and the platform is shown in FIG. 9 qualitatively as a metal temperature behavior which is caused by a thickness difference between thickness of the moving blade trailing edge portion and that of the platform. Likewise, the mentioned relation between the stationary blade trailing edge portion and the shroud is shown in FIG. 10 qualitatively as a metal temperature behavior which is caused by a thickness difference between thickness of the stationary blade trailing edge portion and that of the shroud.
In FIGS. 9 and 10, the vertical axis means a gas turbine rotational speed and metal temperature and the horizontal axis means a lapse of time. When the gas turbine is stopped, gas turbine rotational speed C1, C2 is reduced. In the area of C1 and C2, the blade trailing edge portion which is of a smaller thermal capacity is cooled quicker and moving blade trailing edge portion metal temperature B1 and stationary blade trailing edge portion metal temperature B2 are reduced largely. On the contrary, the platform and the shroud are of a larger thermal capacity, respectively, and platform metal temperature A1 and shroud metal temperature A2 are reduced comparatively slowly. Hence, temperature difference xcex1t between both portions becomes larger and a problem of occurrence of thermal stress arises there.
Thus, in order to solve the problem in the prior art, it is an object of the present invention to provide highly reliable moving blade and stationary blade which are able to suppress an occurrence of thermal stress caused by the mentioned temperature difference as well as to provide a gas turbine equipment comprising these moving blade and stationary blade.
In order to solve the mentioned problem in the prior art, the present invention provides the following first means:
A gas turbine equipment comprising a rotational portion of a rotor and a moving blade, a stationary portion of a casing, a stationary blade, various supporting members and the like and a combustor, characterized in that there is provided a thermal stress reducing portion in any one or both of a moving blade joint adjacent portion between a moving blade trailing edge portion and a platform and a stationary blade joint adjacent portion between a stationary blade trailing edge portion and a shroud.
According to the mentioned first means, the thermal stress reducing portion is provided in any one or both of the moving blade joint adjacent portion between the moving blade trailing edge portion and the platform and the stationary blade joint adjacent portion between the stationary blade trailing edge portion and the shroud, and thereby the undesirable thermal stress is reduced in these joint adjacent portions and the reliability of the gas turbine equipment can be enhanced.
Also, the present invention provides the following second means:
A gas turbine equipment as mentioned in the first means, characterized in that the thermal stress reducing portion provided in the moving blade joint adjacent portion is formed such that the platform in the moving blade joint adjacent portion is partially cut away and a remaining thickness of the platform so cut away is approximately same as a thickness of the moving blade trailing edge portion.
According to the mentioned second means, the thermal stress reducing portion is formed in such a structure that the platform in the moving blade joint adjacent portion between the moving blade trailing edge portion and the platform is partially cut away and a remaining thickness of the platform so cut away is approximately same as a thickness of the moving blade trailing edge portion, and thereby the undesirable thermal stress is surely reduced by the simply workable means and the reliability of the gas turbine equipment can be enhanced.
Also, the present invention provides the following third means:
A gas turbine equipment as mentioned in the first means, characterized in that the thermal stress reducing portion provided in the stationary blade joint adjacent portion is formed such that the shroud in the stationary blade joint adjacent portion is thinned and a remaining thickness of the shroud so thinned is approximately same as a thickness of the stationary blade trailing edge portion.
According to the mentioned third means, the thermal stress reducing portion is formed in such a structure that the shroud in the stationary blade joint adjacent portion between the stationary blade trailing edge portion and the shroud is thinned and a remaining thickness of the shroud so thinned is approximately same as a thickness of the stationary blade trailing edge portion, and thereby the undesirable thermal stress is surely reduced by the simply workable means and the reliability of the gas turbine equipment can be enhanced.
Also, the present invention provides the following fourth means:
A turbine blade comprising a moving blade joint adjacent portion between a moving blade trailing edge portion and a platform, characterized in that the platform in the moving blade joint adjacent portion is partially cut away and a remaining thickness of the platform so cut away is approximately same as a thickness of the moving blade trailing edge portion.
According to the mentioned fourth means, the structure is employed such that the platform in the moving blade joint adjacent portion between the moving blade trailing edge portion and the platform is partially cut away and a remaining thickness of the platform so cut away is approximately same as a thickness of the moving blade trailing edge portion, and thereby the undesirable thermal stress occurring in the moving blade joint adjacent portion is reduced and the reliability of the turbine blade can be enhanced.
Also, the present invention provides the following fifth means:
A turbine blade comprising stationary blade inner and outer joint adjacent portions between a stationary blade trailing edge portion and an inner shroud and between said stationary blade trailing edge portion and an outer shroud, respectively, characterized in that each of the inner shroud in the stationary blade inner joint adjacent portion and the outer shroud in the stationary blade outer joint adjacent portion is thinned and a remaining thickness each of the inner shroud and the outer shroud so thinned is approximately same as a thickness of the stationary blade trailing edge portion.
According to the mentioned fifth means, the structure is employed such that each of the inner shroud in the stationary blade inner joint adjacent portion between the stationary blade trailing edge portion and the inner shroud and the outer shroud in the stationary blade outer joint adjacent portion between the stationary blade trailing edge portion and the outer shroud is thinned and a remaining thickness each of the inner shroud and the outer shroud so thinned is approximately same as a thickness of the stationary blade trailing edge portion, and thereby the undesirable thermal stress occurring in the stationary blade inner and outer joint adjacent portions is reduced and the reliability of the turbine blade can be enhanced.
Further, the present invention provides the following sixth means:
A gas turbine equipment comprising the turbine blade mentioned in the fourth means and that mentioned in the fifth means.
According to the mentioned sixth means, the structure is employed such that, on the moving blade side, the platform in the moving blade joint adjacent portion between the moving blade trailing edge portion and the platform is partially cut away and, on the stationary blade side, each of the inner shroud in the stationary blade inner joint adjacent portion between the stationary blade trailing edge portion and the inner shroud and the outer shroud in the stationary blade outer joint adjacent portion between the stationary blade trailing edge portion and the outer shroud is thinned, and thereby the undesirable thermal stress occurring both on the moving blade side and on the stationary side is reduced and the reliability of the gas turbine equipment can be enhanced.